The present invention generally relates to gas turbine engine systems and, more particularly, to gas turbine cooled shroud assemblies.
Turbine shroud assemblies have been used extensively in gas turbine engines. The turbine shroud assembly may be positioned immediately downstream of a high pressure turbine (HPT) nozzle. The turbine shroud assembly may surround a HPT rotor and may define an outer boundary of a high temperature gas flow path through the HPT. During engine operation, exposure to the high temperature gas flow may result in failure of the turbine shroud components. Due to the differing expansion of rotor and turbine shroud assembly components, it may also result in contact between the turbine shroud assembly and the blade tips of the rotor. A small amount of cooling air from a compressor may be used to decrease some of the adverse effects of the high temperature gas flow.
Minimizing the amount of air necessary to cool the turbine shroud assembly is desirable because engine efficiency decreases as the amount of cooling air increases. Methods for minimizing the cooling air necessary may include decreasing cooling air leakage from the assembly or reducing the cooling needs of the system by increasing the effectiveness of the cooling scheme.
Turbine shroud assemblies have experienced significant distress due to a lack of robust sealing of the assembly. This leakage may result in a significant reduction in the cooling cavity pressure (and back flow margin), which can result in hot gas ingestion and distress in the hardware. Back flow margin is the ratio of the difference between the shroud cooling cavity pressure and the flow path pressure to the flow path pressure. If the back flow margin of the assembly becomes negative (or for some designs even a low positive number), hot flow path gas may ingest into portions of the shroud and can cause significant distress. One challenge in maintaining good back flow margin is due to the difficulty in sealing the various leak paths that allow the cooling air to escape from the shroud cooling cavity.
Several methods of reducing cooling air leakage have been disclosed. These methods include the use of metallic feather type seals and metallic platform seals. Unfortunately, platform seals are not suitable for some applications, and the metallic feather seals, which are secured in machined grooves in the sides of the segments, may fail in the operating environment of some engines. In addition, assembly technicians may cut themselves on the small, sharp metallic platform seals.
Methods of reducing system cooling needs have also been disclosed. Manufacturing the assembly components from more robust materials and utilizing Thermal Barrier Coatings (TBC) have been described. Designs that utilize TBCs to keep the shrouds insulated from the hot flow path gas can experience delamination of the TBC, which in turn results in shroud distress. The shroud distress can result in large turbine blade tip clearances. The subsequent increase in turbine blade tip clearance increases fuel consumption and also results in an increase in turbine inlet temperature, which further distresses the hardware.
Methods of increasing the effectiveness of cooling configurations have been disclosed. In one method complex arrays of film cooling holes have been drilled into shroud segments. Although, this results in increased cooling of the turbine shroud assembly, all edges of the shroud segments may not be sufficiently cooled and system integrity may suffer.
Turbine shroud assemblies having increased cooling of the shroud segment edges have been disclosed. One such disclosure utilizes an interlocking hook/shelf on the ends of the segments in conjunction with conventional feather seals and slots to produce an end gap seal between the adjacent circumferential segments. In addition, this disclosure uses film cooling holes to reinforce cooling at the sides of the segment. Although, cooling of the shroud segment edges is increased, the metallic feather seals may suffer distress at higher operating temperatures due to hot gas ingestion, resulting in a loss of back flow margin to the assembly.
Turbine shroud assemblies having reduced hot gas ingestion have been disclosed in U.S. Pat. No. 4,573,866. These assemblies utilize pin fins to cool the shroud segments and incorporate sheet metal seals and bellows seals to reduce cooling flow leakage. A cooling flow is used to pressurize the area around the shroud segment sides. A feather seal and a tongue-and-groove interlocking feature for adjacent segments are also utilized. Although this results in a reduction of hot gas ingestion at the sides of the segments, hot gas ingestion at the shroud forward and aft cavities may not be sufficiently reduced.
Shroud cooling assemblies having improved cooling of the aft C-clip have been disclosed in U.S. Pat. No. 6,139,257. Cooling holes are formed in the aft rail of the shroud segments to impingement cool the aft corners of the shroud and to pressurize the aft cavity between the base of the shroud segment and the C-clip. Although hot gas ingestion and consequent overheating of the aft corners of the shroud may be reduced, assemblies having further reductions in hot gas ingestion are needed.
As can be seen, there is a need for improved turbine shroud assemblies. Additionally, turbine shroud assemblies are needed wherein hot gas flow ingestion is decreased. Further, assemblies are needed wherein cooling air flow is minimized while allowing for increased gas flow temperatures. Assemblies are needed wherein hot gas ingestion at the shroud forward and aft cavity is reduced. Moreover, turbine shroud assemblies having increased cooling to the high pressure seals and the shroud hangers are needed.